Two stage combustor with reformer

Abstract

The present invention provides a combustor for an aerospace gas turbine engine comprising two stages wherein each stage defines an inlet and an exit. The second stage inlet is in fluid communication with the first stage exit such that a first flowpath is defined and it passes substantially through the second stage. A plurality of flow channel tubes is positioned within the second stage and each flow channel tube passes sealingly through a header plate positioned upstream of the second stage inlet thereby defining a second flowpath that also passes substantially through the second stage. The first flowpath exit and the second flowpath exit are positioned adjacent and proximate to one another to provide for the generation of microflames or microflame jets exiting the second stage from between and around the flow channel tube exits. The first stage of the combustor provides a gasifier and a reformer. The present invention also may comprise an igniter for further combustion of the reacted products or an external heat source for start-up. The second stage also may comprise a microflame combustor.

The invention claimed is: 
     
       1. A combustor for an aerospace gas turbine engine comprising:
 a) a first stage wherein the first stage defines a reformer comprising a first stage inlet and a first stage exit, the first stage inlet adapted to input all of a fuel to the reformer; the reformer further comprising a metal screen substrate and adapted to dry reform a fuel-rich mixture of fuel and air into a reacted fuel stream comprising carbon monoxide and hydrogen; 
 b) a second stage wherein the second stage defines a second stage inlet and a second stage exit, and wherein the second stage inlet is in fluid communication with the first stage exit thereby defining a first flowpath having a first flowpath inlet, and wherein the first flowpath passes substantially through the second stage and defines a first flowpath exit; 
 c) a plurality of flow channel tubes positioned within the second stage wherein each flow channel tube defines a flow channel tube inlet and a flow channel tube exit, and wherein each flow channel tube inlet passes sealingly through a header plate positioned upstream of the second stage inlet thereby defining a second flowpath having a second flowpath inlet defined by the plurality of flow channel tube inlets, and wherein the second flowpath passes substantially through the second stage and defines a second flowpath exit defined by the plurality of flow channel tube exits, further wherein the first flowpath adapted to pass the reacted fuel stream comprising carbon monoxide and hydrogen within the second stage is configured to be in heat exchange with the second flowpath adapted to pass combustion air; and 
 (d) wherein the first flowpath exit and the second flowpath exit, defined at the termination of the plurality of flow channel tubes, are positioned adjacent and proximate to one another. 
 
     
     
       2. The combustor for the aerospace gas turbine of  claim 1  wherein the first stage comprises a gasifier. 
     
     
       3. The combustor for the aerospace gas turbine of  claim 1  wherein the reformer further comprises an ultra-short-channel-length substrate. 
     
     
       4. The combustor for the aerospace gas turbine of  claim 3  wherein the ultra-short-channel-length substrate reformer provides conversion of jet fuel to a reacted fuel stream comprising carbon monoxide and hydrogen. 
     
     
       5. The combustor for the aerospace gas turbine of  claim 1  further comprising an igniter. 
     
     
       6. The combustor for the aerospace gas turbine of  claim 1  further comprising an external heat source for start-up. 
     
     
       7. The combustor for the aerospace gas turbine of  claim 1  wherein the second stage further comprises a microflame combustor. 
     
     
       8. The combustor of  claim 1  further configured to provide for 30 percent of the total air flow to be passed into the reformer. 
     
     
       9. The combustor of  claim 1  further configured to provide for less than 5 percent unmixedness between the heat exchanged reaction mixture and the combustion air at a distance within 1 to 2 inches from the exits of the first and second flowpaths. 
     
     
       10. The combustor of  claim 1  wherein the first stage reformer is positioned off center with respect to a centerline of the second stage of the combustor. 
     
     
       11. A method for combusting jet fuel comprising:
 a) reacting a fuel-rich mixture of jet fuel and air in a first stage of a combustor, the first stage comprising a reformer adapted to dry reform a mixture comprising air and jet fuel to form a reacted mixture comprising carbon monoxide and hydrogen, further wherein the reformer comprises a metal screen substrate and the reformer is adapted to input all of the jet fuel; 
 b) passing the reacted mixture comprising carbon monoxide and hydrogen into a first flowpath within a second stage of the combustor, the second stage being adapted to provide heat exchange with combustion air to form a heat exchanged mixture comprising carbon monoxide and hydrogen; the combustion air being provided within a second flowpath within the second stage of the combustor, the second flowpath comprising a plurality of flow channel tubes through which the combustion air passes in heat exchange with the first flowpath; and 
 c) passing the heat exchanged mixture of carbon monoxide and hydrogen, exiting the first flowpath as microjets around and between the plurality of flow channel tubes, into contact with the combustion air exiting the second flowpath for combustion. 
 
     
     
       12. The method of  claim 11  wherein 30 percent of total air flow is passed into the reformer. 
     
     
       13. The method of  claim 11  wherein fuel conversion in the reformer ranges from 20 to 30 percent.